The present invention relates to a panel adapted to control aerodynamic phenomena on a body immersed in a subsonic fluid current.
Aerodynamic phenomena occurring in transonic conditions on a body having a conical, cylindrical or another aerodynamically “squat” shape create shock waves that interact with the development of the boundary layer, thus causing disorderly recirculation phenomena or air flow turbulences.
Such turbulences may consist, for example, of alternated separation and reattachment of vortex structures; this phenomenon is called vortex shedding.
A body comprising at least one circular or semicircular or elliptical or aerodynamically squat portion creates a trail of vortices detaching from the body itself in an alternated, non-stationary manner when the body itself is hit by a subsonic fluid flow.
The detachment frequency and intensity of the vortices are a function of the size of the body itself and of the incident fluid current. This phenomenon increases aerodynamic resistance, leading to the structure being subject to greater stress. Said stress involves both the body itself and the aircraft portion whereto said body has been applied.
Such aerodynamic phenomena, like vortex shedding, arise on high-performance civil and military aircraft. Of course, aircraft of this type include a computerized flight control system, abbreviated with the acronym FCS.
To be fully functional, such a system requires real-time acquisition of a plurality of parameters, such as, for example, aircraft position and trim and flight conditions in terms of speed, height and pressure. These parameters are normally referred to as air data, the acquisition of which is entrusted to an additional data acquisition system called air data sensor or ADS.
A typical system for acquiring said parameters performs multiple acquisition of redundant data.
The redundant acquisition of said parameters requires the installation, in various parts of the aircraft, of at least two structures, within which a plurality of sensors and/or probes are suitably arranged in order to acquire said air data.
Such a structure is typically cylindrical or conical in shape, referred to as integrated multi function probe or IMFP. Said IMFP has a substantially longitudinal extension, perpendicular to the flight direction and normal to the surface of the aircraft.
Through said sensors and/or probes, which are arranged in a plurality of slots provided at different angles of the circular section of said IMFP, the acquisition system can acquire the air data, such as local pressure.
The shape of said circular-section IMFP's causes aerodynamic phenomena, like vortex shedding, when subjected to a subsonic flow.
Vortex shedding causes a pressure drop downstream of said IMFP, in the flow direction, resulting in erroneous detection of data which are essential to the aircraft's computerized control system.
Furthermore, in particular flight and height conditions the vortex shedding phenomenon also causes increased noise in frequency bands which can also be heard by the human ear.
In summary, the most important problems caused by these aerodynamic phenomena are:                structural fatigue caused by increased aerodynamic resistance;        noise generation;        acquisition of disturbed signals, e.g. due to local pressure drop.        
The prior art has attempted to solve these technical problems by acting upon the shape of said body or IMFP, by adapting its profile in order to make it more aerodynamic and minimize the vortex shedding phenomenon.
Sensors, antennae or aerodynamic excrescences are typically installed to the outer covering of an airplane on a smooth surface created with special care from the aerodynamic viewpoint, so as to minimize wall-triggered turbulences that might hit the body and jeopardize the proper operation thereof. Directly modifying the shape of the body or object in order to make it compatible with the local aerodynamic flow field is not always possible.
The shape of the body is often dictated by functional requirements. Functional requirements often deviate from aerodynamic ones, so that the body will tend to develop aerodynamic phenomena, such as turbulent and vortical structures, which may originate the above-described problems.
These solutions cannot effectively eliminate such aerodynamic phenomena because aerodynamic phenomena vary as a function of many parameters, so that it is difficult to create an aerodynamic profile capable of ensuring high-performance in every flight condition of an aircraft, such as speed, altitude, etc.
The solutions known in the art, in fact, are aimed at solving particular aerodynamic phenomena in particular flight conditions, but they may become sources of other uncontrollable aerodynamic phenomena in other flight conditions, different from those for which they have been designed or studied.
U.S. Pat. No. 4,664,345 has disclosed a device that can eliminate the problems relating to the separation of the boundary layer of a downstream laminar flow, with respect to the air flow, caused by a step in the aircraft structure. Said device is unidirectional.
It is also know from the above-mentioned US patent that the device has 2 distinct chambers connected by a valve, each chamber being assigned a univocal function as flow inlet or outlet.
It is known from U.S. Pat. No. 4,664,345 a device comprising suction inlets in the surface just upstream of a disturbance and blowing outlets just downstream of the disturbance and by a flow channel interconnecting these inlets and outlets. Passage of a portion of the flowing medium through these passages is automatically assured due to a pressure differential between the inlets and outlets.
Furthermore it is also known from U.S. Pat. No. 5,628,565 an aerodynamic air data sensing probe adapted for mounting to an air vehicle and capable of generating signals related to a fluid flowing relative to the air vehicle. A fluid inlet positioned on a first end of the strut faces generally transverse to the fluid flow selectively admits fluid to an internal strut cavity due to a pressure differential there across. In operation, the pressure differential forms between the first end surface of the aerodynamically-shaped, forward-inclined strut and probe exhaust ports.